FIG. 1 illustrates a gas turbine aircraft engine as known in the art. Hot gases 3 provided by a combustor 6 impart energy to a high pressure turbine 9 which is surrounded by a shroud 12. A tip clearance 15 exists between the turbine 9 and the shroud 12. When the engine is idling, say at 11770 rpm, a given tip clearance will exist, such as 0.047 inches. However, during a sudden acceleration, such as to 16140 rpm within 4 seconds, increases in centrifugal forces due to the greater rotational speed causes the turbine to expand, thus reducing the tip clearance to 0.016 inches. Temperature in the region of the high pressure turbine 9 will increase, thus causing the turbine 9 and the shroud 12 to expand thermally. However, since the thermal mass of the shroud 12 is much less than that of the turbine 9, the shroud 12 heats up faster, and thus expands faster. Accordingly, the tip clearance 15 initially increases, say to 0.028 inches, and then decreases to the steady-state level of 0.016 inches when the temperature of the turbine 9 reaches its steady-state value.
This initial increased tip clearance 15 is undesirable because it imposes a penalty in engine efficiency: the hot gases 3 can bypass the turbine 9 by leaking through the tip clearance region 15, and thus the leaking gases do little or no work upon the turbine 9.